The present invention relates to a tip turbine engine, and more particularly to an assembly with a load bearing support shaft cantilevered from a single engine support plane.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and a low pressure compressor, a middle core engine including a combustor, and an aft low pressure turbine all located along a common longitudinal axis. Although highly efficient, conventional turbofan engines operate in a axial flow relationship. The serial flow relationship results in a relatively complicated elongated engine structure and, therefore, requires multiple mounting planes to mount the engine and bear the loads of the elongated structure. Utilizing multiple mounting planes may complicate the mounting process and make assembly laborious and expensive.
A recent development in gas turbine engines is the more longitudinally compact tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan, which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor located radially outward from the fan. The combustor ignites the fuel mixture to form a high energy gas stream which drives turbine blades that are integrated onto the tips of the hollow bypass fan blades for rotation therewith as disclosed in U.S. Patent Application Publication Nos.: 2003192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust to weight ratio equivalent to conventional turbofan engines of the same class within a package of significantly shorter longitudinal length.
Accordingly and because of the unique architecture and shorter length of the tip turbine engine, it is desirable to cantilever the length of the engine from a sole engine support plane.